Electric-based secondary power system architectures for aircraft

ABSTRACT

Methods and systems for providing secondary power to aircraft systems. In one embodiment, an aircraft system architecture for providing power to an environmental control system includes an electric generator operably coupled to a jet engine. The jet engine can be configured to provide propulsive thrust to the aircraft, and the electric generator can be configured to receive shaft power from the jet engine. The environmental control system can be configured to provide outside air to a passenger cabin of the aircraft in the absence of bleed air from the jet engine.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application is a continuation of U.S. application Ser. No.10/691,440, filed Oct. 21, 2003, and this application claims priority toU.S. Provisional Patent Application Ser. No. 60/420,637, filed Oct. 22,2002 and incorporated herein in its entirety by reference. Thisapplication incorporates U.S. Pat. No. 6,526,775 in its entirety byreference, and this application relates to U.S. application Ser. No.11/301,179, filed Dec. 12, 2005.

TECHNICAL FIELD

The following disclosure relates generally to secondary power systemsfor aircraft and, more particularly, to electric-based secondary powersystems for aircraft.

BACKGROUND

Conventional transport aircraft typically utilize pneumatic, hydraulic,and electric power from main engines to support various aircraft systemsduring flight. In addition, conventional transport aircraft typicallyutilize pneumatic and electric power from on-board auxiliary power units(APUs) to support aircraft systems during ground operations. Aircraftair conditioning systems are typically the largest secondary power userson commercial transport aircraft. On conventional transport aircraft,these systems use high temperature/high pressure air extracted from theengine compressor stages (“bleed air”). The air passes through airconditioning packs before passing into the fuselage to meet temperature,ventilation, and pressurization needs. The conditioned air is thendischarged from the fuselage through outflow valves or through normalcabin leakage. During ground operations, the APU can provide bleed aireither from a separate shaft-driven load compressor or from a powersection compressor. Similar to the bleed air from the main engines, thehigh temperature and high pressure air from the APU passes through airconditioning packs before passing into the fuselage.

FIG. 1 schematically illustrates a conventional pneumatic-basedsecondary power system architecture 100 configured in accordance withthe prior art. The system architecture 100 can include jet engines 110(shown as a first engine 110 a and a second engine 110 b) for providingpropulsive thrust to the aircraft (not shown). In addition to thrust,the engines 110 can also provide high temperature/high pressure air to ableed manifold 120 via bleed ports 112 (identified individually as afirst bleed port 112 a and a second bleed port 112 b). The bleed ports112 receive air from the compressor stages of the engines 110, and passthe air through heat exchangers 114 (such as precoolers) that cool theair before it passes to the bleed manifold 120.

The high pressure air from the bleed manifold 120 supports the majorityof secondary power needs of the aircraft. For example, a portion of thisair flows to air conditioning packs 140 (shown as a first airconditioning pack 140 a and a second air conditioning pack 140 b) thatsupply conditioned air to a passenger cabin 102 in a fuselage 104. Theair conditioning packs 140 include a series of heat exchangers,modulating valves, and air cycle machines that condition the air to meetthe temperature, ventilation, and pressurization needs of the passengercabin 102. Another portion of air from the bleed manifold 120 flows toturbines 160 that drive high capacity hydraulic pumps 168. The hydraulicpumps 168 provide hydraulic power to the landing gear and otherhydraulic systems of the aircraft. Yet other portions of this highpressure air are directed to an engine cowl ice protection system 152and a wing ice protection system 150.

The wing ice protection system 150 includes a valve (not shown) thatcontrols the flow of bleed air to the wing leading edge, and a “piccolo”duct (also not shown) that distributes the hot air evenly along theprotected area of the wing leading edge. If ice protection of leadingedge slats is required, a telescoping duct can be used to supply hotbleed air to the slats in the extended position. The ice protectionbleed air is exhausted through holes in the lower surface of the wing orslat.

In addition to the engines 110, the system architecture 100 can alsoinclude an APU 130 as an alternate power source. The APU 130 istypically started by a DC starter motor 134 using a battery 136. The APU130 drives a compressor 138 that provides high pressure air to the bleedmanifold 120 for engine starting and other ground operations. For enginestarting, the high pressure air flows from the bleed manifold 120 tostart-turbines 154 operably coupled to each of the engines 110. As analternative to the APU 130, bleed air from a running one of the engines110 can be used to re-start the other engine 110. As a furtheralternative, an external air cart (not shown in FIG. 1) can provide highpressure air for engine starting on the ground.

The system architecture 100 can further include engine-driven generators116 operably coupled to the engines 110, and an APU-driven generator 132operably coupled to the APU 130. In flight, the engine-driven generators116 can support conventional electrical system loads such as a fuel pump108, motor-driven hydraulic pumps 178, and various fans, galley systems,in-flight entertainment systems, lighting systems, avionics systems, andthe like. The APU-driven generator 132 can support these functionsduring ground operations and during flight as required. Theengine-driven generator 116 and the APU-driven generator 132 aretypically rated at 90-120 kVA and produce a voltage of 115 Vac. They canprovide power to transformer-rectifier units that convert 115 Vac to 28Vdc for many of the abovementioned electrical loads. The power isdistributed through an electrical system based largely on thermalcircuit-breakers and relays.

The system architecture 100 can additionally include engine-drivenhydraulic pumps 118 operably coupled to the engines 110. The hydraulicpumps 118 provide hydraulic power to control surface actuators and otheraircraft systems in flight. Electric-motor driven pumps 178 can provideback-up hydraulic power for maintenance activities on the ground.

FIG. 2 is a schematic top view of a prior art aircraft 202 that includesthe secondary power system architecture 100 of FIG. 1. The aircraft 202includes a forward electronic equipment bay 210 that distributeselectrical power to a plurality of electrical loads 220 associated withthe system architecture 100 described above. In flight, the electronicequipment bay 210 can receive electrical power from the enginegenerators 116, as well as the APU 130. On the ground, the electronicequipment bay 210 can receive electrical power from the APU 130, or froman external power source 212 via a receptacle 213.

One shortcoming of the secondary power system architecture 100 describedabove is that it is sized for a worst case operating condition(typically, cruise speed, high aircraft load, hot day, and one enginebleed air system failed) to ensure sufficient air flow is available tomeet system demands at all times. As a result, under typical operatingconditions, the engines 110 provide bleed air at a significantly higherpressure and temperature than the air conditioning packs 140 and theother aircraft systems demand. To compensate, the precoolers 114 and theair conditioning packs 140 regulate the pressure and temperature tolower values as required to meet the demands for fuselagepressurization, ventilation, and temperature control. Consequently, asignificant amount of energy is wasted by precoolers and modulatingvalves during this regulation. Even under optimum conditions, asignificant amount of energy extracted from the engines 110 is wasted inthe form of heat and pressure drops that occur in the ducting, valvesand other components associated with the bleed manifold 120 and the airconditioning packs 140.

SUMMARY

The present invention is directed generally toward secondary powersystems for aircraft and methods for providing secondary power toaircraft systems. In one embodiment, an aircraft configured inaccordance with one aspect of the invention includes a fuselage and ajet engine configured to provide propulsive thrust to the aircraft. Theaircraft can further include an electric generator operably coupled tothe jet engine, and an environmental control system. The environmentalcontrol system can include at least one compressor motor configured toreceive electric power from the electric generator to provide outsideair to the fuselage in the absence of bleed air from the jet engine.

In another aspect of this embodiment, the aircraft can include a wingextending outwardly from the fuselage, and an electrothermal wing iceprotection system. The electrothermal wing ice protection system can beconfigured to receive electric power from the electric generator to atleast reduce the formation of ice on a portion of the wing in theabsence of bleed air from the jet engine. In a further aspect of thisembodiment, the electric generator can be a first electric generator,and the aircraft can additionally include an auxiliary power unit and asecond electric generator. The second electric generator can be operablycoupled to the auxiliary power unit and configured to receive shaftpower from the auxiliary power unit. In this aspect, the at least onecompressor motor of the environmental control system can be configuredto receive electric power from the second electric generator to provideoutside air to the passenger cabin in the absence of compressed air fromthe auxiliary power unit.

In another embodiment, a method for providing conditioned air to afuselage of an aircraft can include providing a compressor fan in flowcommunication with the fuselage, and operably coupling an electric motorto the compressor fan to drive the compressor fan. The method canfurther include operably coupling an electric generator to a jet engineof the aircraft, and providing electric power from the electricgenerator to the electric motor to drive the compressor fan. In oneaspect of this embodiment, the compressor fan can be driven to flow airfrom outside the fuselage into the fuselage in the absence of bleed airfrom the jet engine.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic diagram of a conventional pneumatic-basedsecondary power system architecture configured in accordance with theprior art.

FIG. 2 is a schematic top view of a conventional aircraft having thesystem architecture of FIG. 1.

FIG. 3 is a schematic diagram illustrating an electric-based secondarypower system architecture configured in accordance with an embodiment ofthe invention.

FIG. 4 is a schematic top view of an aircraft having the systemarchitecture of FIG. 3 configured in accordance with an embodiment ofthe invention.

FIG. 5 is a schematic diagram of an aircraft electric power distributionsystem configured in accordance with an embodiment of the invention.

FIG. 6 is a schematic diagram of an aircraft electric power distributionsystem configured in accordance with another embodiment of theinvention.

FIG. 7 is a schematic diagram of an aircraft electric power distributionsystem having only AC generators in accordance with a further embodimentof the invention.

FIGS. 8A-8C are schematic diagrams illustrating an electric powerdistribution system having an engine start circuit configured inaccordance with an embodiment of the invention.

DETAILED DESCRIPTION

The following disclosure describes systems and methods for providingpower to aircraft systems. Certain details are set forth in thefollowing description and in FIGS. 3-7 to provide a thoroughunderstanding of various embodiments of the invention. Other detailsdescribing well-known structures and systems often associated with theaircraft and/or aircraft secondary power systems are not set forth inthe following disclosure to avoid unnecessarily obscuring thedescription of the various embodiments of the invention.

Many of the details, dimensions, angles, and other features shown in theFigures are merely illustrative of particular embodiments of theinvention. Accordingly, other embodiments can have other details,dimensions, and features without departing from the spirit or scope ofthe present invention. In addition, further embodiments of the inventionmay be practiced without several of the details described below.

In the Figures, identical reference numbers identify identical or atleast generally similar elements. To facilitate the discussion of anyparticular element, the most significant digit or digits of anyreference number refer to the Figure in which that element is firstintroduced. For example, element 310 is first introduced and discussedwith reference to FIG. 3.

FIG. 3 is a schematic diagram illustrating an electric-based secondarypower system architecture 300 configured in accordance with anembodiment of the invention. In one aspect of this embodiment, thesystem architecture 300 includes a first engine 310 a and a secondengine 310 b for providing propulsive thrust to an aircraft (not shown).As described in greater detail below, a first starter/generator 316 aand a second starter/generator 316 b can be operably coupled to each ofthe engines 310 to provide electrical power to a plurality of aircraftsystems on an on-demand basis. The starter/generators 316 support amajority of the aircraft functions that were traditionally performed bythe bleed-air system described above in FIG. 1. These functions caninclude fuselage air conditioning and pressurization, engine starting,and wing ice protection among others.

In another aspect of this embodiment, the system architecture 300further includes an APU 330 for providing power to aircraft systems whenneeded during ground operations and in flight. Power for starting theAPU 330 can be provided by an aircraft battery 336, an external groundpower source (not shown), or one or more of the engine-drivenstarter/generators 316. Power from the APU 330 is provided by a firstAPU starter/generator 332 a and a second APU starter/generator 332 b,each of which are operably coupled to the APU 330.

In contrast to the conventional APU 130 described above with referenceto FIG. 1, the APU 330 provides only electric power to the variousaircraft systems. Consequently, it can be much simpler than the APU 130because all of the components associated with pneumatic power deliverycan be eliminated. This feature can result in a significant improvementin APU reliability and a reduction in required maintenance.

In a further aspect of this embodiment, the system architecture 300includes an environmental control system having a first air conditioningpack 340 a and a second air conditioning pack 340 b. The airconditioning packs 340 are configured to provide conditioned air to apassenger cabin 302 in a fuselage 304 to meet temperature, pressure, andair conditioning needs. In one embodiment, the air conditioning packs340 can be at least generally similar to one or more of the airconditioning systems disclosed in U.S. Pat. No. 6,526,775, which isincorporated herein in its entirety by reference. In another embodiment,the air conditioning packs 340 can include adjustable speed electriccompressor motors 380 configured to receive electric power from theengines 310 during flight and the APU 330 during ground operations. Thecompressor motors 380 drive compressors (not shown) that receive freshoutside air via ram air inlets 342. The fresh air is compressed andflows from the air conditioning packs 340 into the fuselage 304 to meetthe pressurization and temperature control needs of the cabin 302. Inone embodiment, the system architecture 300 can include one or morevariable speed fans (not shown) to distribute the air to various partsof the fuselage 304 at different flow rates to meet the particulardemands of the fuselage 304 at any given time. Tailoring the power drawfrom the engines 310 in this manner can further increase fuelefficiency.

The adjustable speed compressor motors 380 allow the cabin air pressureand air flow to be varied based on cabin volume, occupant count, and/orthe desired cabin pressure altitude. For example, if a lower cabinaltitude is desired (higher pressure), then the electric ECS system ofthe present invention can accommodate this by increasing inflow with theadjustable speed compressor motors 380 and/or decreasing outflow fromthe fuselage 304. In general, conventional pneumatic systems do not havethe ability to lower cabin altitudes much below their design points(e.g., 8000 ft) because the systems are typically sized for the designpoint. Another benefit of the electric approach to air conditioning overthe conventional pneumatic approach is that the energy extracted fromthe engines for the electric approach is not wasted by pre-coolers andmodulating valves in the air conditioning packs 340. Instead, thecompressor motors 380 only draw enough electric power from the engines310 as is required by the adjustable speed compressors to meet theimmediate pressurization needs of the cabin 302. This real-time energyoptimization can be extended to other electric power users across theaircraft platform to improve fuel efficiency. As described below, forexample, such users can include recirculation fans, Lavatory and galleyvent fans, cargo heating, wing ice protection, and hydraulic actuation.By only drawing the energy needed, fuel economy can be increased.

In another aspect of the invention, the system architecture 300 furtherincludes a wing ice protection system 350 that utilizes electrical powerfrom the engines 310. The wing ice protection system 350 can beconfigured in accordance with at least two embodiments of the presentinvention to prevent or at least reduce the formation of ice on aportion of a wing 352. In an electrothermal ice protection embodiment,heating elements such as blankets (not shown) can be bonded or otherwisepositioned proximate to interior portions of the wing leading edges. Forwing ice protection, the heating blankets can be energized sequentiallyto heat the wing leading edge causing any ice build-up to melt and/ordetach from the wing leading edge. This method can be significantly moreefficient than conventional bleed air systems because desired portionsof the wing leading edge are heated sequentially rather thansimultaneously. Consequently, the power draw for ice protection issignificantly reduced. In addition, in contrast to bleed air systems,there are no bleed air exhaust holes on the wings. As a result, aircraftdrag and community noise are reduced relative to conventional systems.

The wing ice protection system 350 can also operate as anelectromechanical system in accordance with another embodimentinvention. In this embodiment, electromechanical actuators (not shown)in an interior portion of the wing leading edges can be configured tobriefly vibrate the wing leading edge, causing any ice build-up todetach and fall away. This embodiment may require significantly lesselectrical power than the electrothermal embodiment discussed above. Ineither embodiment, the wing ice protection system 350 can be broken upinto different segments that apply to different regions of the wing orslat leading edge. That way, if one portion of the wing leading edgedoes not require ice protection, then that section of the wing iceprotection system 350 can be turned off, resulting in a furtherreduction in power demand from the engines. Additionally, differentsections of the ice protection system can be cycled according todifferent schedules as required to sufficiently reduce ice whileoptimizing power usage.

In another aspect of this embodiment, the starter/generators 316 can bedual-function devices that provide electrical power for aircraft systemswhen operating as generators, and shaft power for engine starting whenoperating as starters. This electrical start capability can enhance thein-flight starting sequence of the engines 310 in the event one or moreof the main engines 310 shuts down during normal flight operations. Forexample, typical high bypass ratio engines may have difficultyrestarting during all flight regimes because the in-flight windmilleffect may not provide enough torque. In contrast, thestarter/generators 316 of the present invention are configured toreceive electric power from any number of electrical sources on theaircraft to assist the engines 310 during an in-flight restart byproviding additional starting torque.

To start the engines 310, the starter/generators 316 can be run assynchronous starting motors with the starting process controlled byengine start converters (not shown). The engine start converters canprovide conditioned electrical power (e.g., adjustable voltage andfrequency) to the starter/generators 316 during the start process foroptimum start performance. The engine start converters can also functionas motor controllers for the cabin pressurization compressor motors 380and/or other adjustable speed motors on the aircraft. Similarly, an APUstart converter (not shown) can function as a motor controller for otheradjustable speed motors on the aircraft such as an on-board inert gasgeneration system (OBIGGS) 309. The power necessary to energize thestarter/generators 316 for engine starting can come from the aircraftbattery 336, the APU 330, a ram air turbine (RAT) 367, a fuel cell (notshown), or other sources. The dual-function aspect of thestarter/generators 316 is not offered by the air turbine engine-starters154 described above with reference to FIG. 1. Unlike thestarter/generators 316, the air turbine engine-starters 154 serve nopurpose while the engines 110 are running.

In a further aspect of this embodiment, the starter generators 316 canbe directly coupled to gear boxes of the engines 310 such that theyoperate at frequencies (e.g., 360-700 Hz) proportional to the enginespeeds. This type of generator may be the simplest and most efficientapproach because the generator does not include a complex constant speeddrive. As a result, in this embodiment the starter/generators 316 may bemore reliable and have lower spares costs than conventional generatorshaving complex constant speed drives. In other embodiments, however,other types of generators can be used. For example, in one otherembodiment where a constant speed is desirable, a constant speedgenerator can be used.

In a further aspect of this embodiment, the system architecture 300includes a hydraulic system that has left, right, and center channels.The hydraulic power for the left and right channels can be provided byengine-driven hydraulic pumps 318 that are operably coupled to each ofthe engines 310. In addition, smaller electric motor-driven hydraulicpumps 319 can also provide hydraulic power to the left and rightchannels for ground operations and to supplement the engine-driven pumps318. The engine-driven pumps 318 can provide hydraulic power for flightcontrol actuators, stabilizer trim actuators, and other functions. Thehydraulic power for the center channel is provided by two large-capacityelectric motor-driven hydraulic pumps 368. In contrast to the hydraulicpumps 168 described above with reference to FIG. 1, which are driven byengine bleed air to meet peak hydraulic demands, the hydraulic pumps 368are driven by electric power from the engines 310. The hydraulic pumps368 can provide hydraulic power for a landing gear system 369 and othersystems, including flight control actuation, thrust reversers, brakes,leading/trailing edge flaps, and a nose gear steering system (notshown). In a further aspect of this embodiment, only one of thehydraulic pumps 368 runs throughout an entire flight, while the otherpump only operates during takeoff and landing.

In another aspect of this embodiment, the system architecture 300 caninclude a plurality of adjustable-speed fuel pumps 308 to transfer fuelfrom a fuel tank 390 to one or more of the engines 310 or to anotherfuel tank (not shown). Typical commercial aircraft use fuel pumps totransfer fuel from one area of the wing to another. This allows theaircraft to maintain a center of gravity that maximizes aircraftperformance. In the conventional pneumatic-based system architecture 100described above with reference to FIG. 1, constant-speed fuel pumps aretypically included for transferring fuel from one tank to the next or tothe engines 110. These constant-speed fuel pumps are typicallyconfigured to operate at a maximum pressure at all times, even thoughthe flow rate corresponding to this maximum pressure is seldom requiredto adequately transfer fuel between tanks or to the engines 1 10. Forthis reason, such fuel systems typically include pressure regulatorsthat simply bleed off the excess fuel pressure. This excess fuelpressure corresponds to wasted engine power. In contrast, in theelectric-based architecture 300 of FIG. 3, the fuel pump speeds can bevaried to transfer fuel from one tank to the next based on the amount offuel needed for transfer and the rate at which the transfer needs tooccur to optimize the overall center of gravity of the aircraft duringnormal flight conditions. The ability to maintain an optimum center ofgravity throughout the flight segment in this manner can further improveaircraft range and fuel efficiency by only drawing the amount of poweractually needed by the fuel pump at any given time.

A number of other systems can be incorporated into the systemarchitecture 300 to further reduce the power extracted from the engines310. For example, in one embodiment, adjustable- or variable-speed fanscan be used in the air conditioning packs 340 that tailor the powerextracted from the engines 310 based on fan speed. In anotherembodiment, resistive heaters can be used to warm cargo holds (notshown) instead of the bleed air used in conventional systems. Theseresistive heaters can have the ability to be pulse-width modulated tobetter control temperature and further reduce energy consumption.Similarly, the cargo air conditioning systems can be configured to relyless on outside air and more on recirculated air for compartmentcooling. In this manner, the energy losses associated with outside airare eliminated and only the power required to cool the recirculated airis expended.

As discussed above with reference to FIG. 1, in the conventional systemarchitecture 100, the engines 110 provide the majority of secondaryaircraft power in pneumatic form from bleed air. In contrast, in thesystem architecture 300 of the present invention, the engines 310provide the majority of secondary aircraft power in electrical form fromthe starter/generators 316. Eliminating the pneumatic bleed ports fromthe compressor portions of the engines 310 results in a more efficientengine design by reducing compressor capacity requirements and improvingthe operating cycle. Furthermore, eliminating the maintenance intensivebleed system is expected to reduce aircraft maintenance needs andimprove aircraft reliability because there are fewer components on theengine and fewer pneumatic ducts, pre-coolers and valves in thedistribution system. In addition, measures to protect against duct burstand over-temperature conditions are unnecessary with the electric-basedsystem architecture 300.

A further advantage of the electric-based system architecture 300 isthat it can utilize motor controllers to tailor the individual loads toextract only the minimum amount of power necessary from the engines 310at any operating condition. Because these loads are adjustable ratherthan simply on or off, less power is withdrawn from the engines 310. Theability to tailor the power consumption for any electrical power loadcan directly improve aircraft fuel efficiency. Other benefits associatedwith the electric-based system architecture 300 can include thefollowing: real-time power extraction optimization and elimination ofwaste associated with engine bleed air; enhanced air quality; potentialreduction in non-recurring engineering associated with certification ofa multiple engine bleed air system.

Although the system architecture 300 described above with reference toFIG. 3 includes two engines with two starter/generators per engine, inother embodiments, system architectures configured in accordance withthe present invention can include more or fewer engines having more orfewer generators depending on the needs of the particular application.For example, in one other embodiment, a system architecture configuredin accordance with the present inventing can include four jet engines,each having a single starter/generator. In yet another embodiment, asystem architecture configured in accordance with the present inventioncan include only a single engine having two or more generators orstarter/generators. Accordingly, the invention is not limited toaircraft having a particular number of engines or starter/generators.

In addition, although the engines 310 described above with reference toFIG. 3 provide electric power for secondary aircraft systems via thestarter/generators 316, in other embodiments, the engines 310 can alsoinclude one or more bleed ports similar to those disclosed in FIG. 1 forproviding pneumatic power to one or more secondary systems. In theseembodiments, power can be provided to one or more of the airconditioning packs 340, the wing ice protection system 350, and/or thehydraulic pump 368 in the form of pneumatic power or pneumatic andelectric power. Similarly, in other embodiments, the APU 330 can alsoprovide pneumatic power in addition to electric power. Accordingly, theinvention is not limited to aircraft utilizing strictly electric powerfor secondary systems, but can extend in various embodiments to aircraftusing various combinations of electric power, pneumatic power, orelectric and pneumatic power.

FIG. 4 is a schematic top view of an aircraft 402 that includes theelectric-based system architecture 300 of FIG. 3 configured inaccordance with an embodiment of the invention. The aircraft 402 caninclude the four starter/generators 316 coupled to the two engines 310,and the two starter/generators 332 coupled to the APU 330 mounted in thetail of the aircraft 402. In one aspect of this embodiment, the aircraft402 can further include two ground receptacles 413 (identified as afirst ground receptacle 413 a and a second ground receptacle 413 b)configured to receive 115 Vac or 230 Vac power from external powersources 412 a and 412 b, respectively.

In another aspect of this embodiment, the aircraft 402 can include aforward electrical equipment bay 410 a and an aft electrical equipmentbay 410 b. Four remote power distribution units (RPDUs) 424 a-d candistribute electrical power from the equipment bays 410 to a pluralityof system loads 420 associated with the system architecture 300. TheRPDUs 424 can be largely based on solid state power controllers insteadof traditional thermal circuit-breakers and relays.

FIG. 5 is a schematic diagram of an aircraft electric power distributionsystem 500 configured in accordance with an embodiment of the invention.In one aspect of this embodiment, the power distribution system 500includes a first generator 516 a and a second generator 516 b operablycoupled to an aircraft engine 510. In one embodiment, the firstgenerator 516 a and the second generator 516 b can be high voltage ACgenerators (such as 230 Vac generators). In another embodiment, one ofthe two generators 516 can be a high voltage DC generator (such as a±270 Vdc generator). The AC generator 516 a can provide electrical powerto aircraft equipment that is insensitive to the supply frequency. TheDC generator 516 b can provide electrical power to those components ofthe aircraft system that include adjustable speed motors. In otherembodiments, the generators 516 can be other types of generators. Forexample, in one other embodiment, both of the generators 516 can be ACgenerators. In this embodiment, the DC power needs of the system can bemet with suitable AC-to-DC conversion devices. In another embodiment,both of the generators 516 can be DC generators, and the AC power needsof the system can be met with suitable DC-to-AC conversion devices.

In another aspect of this embodiment, the power distribution system 500can further include a first bus 515 a configured to receive power fromthe first generator 516 a, and a second bus 515 b configured to receivepower from the second generator 516 b. In one embodiment, the first bus515 a can be a high voltage AC bus, such as a 230 Vac bus, configured tosupply power directly to a plurality of large-rated AC loads 550. Suchloads may be associated with wing ice protection equipment, hydraulicpumps, fuel pumps, galley systems, and the like. In addition, the firstbus 515 a can also provide power directly to a third bus 515 c via astep-down transformer 522. In one embodiment, the third bus 515 c can bea lower voltage AC bus, such as a 115 Vac bus. The third bus 515 c canprovide power to a plurality of small-rated AC equipment loads 544 onthe aircraft via a plurality of RPDUs (identified as a first RPDU 524 aand at least a second RPDU 524 b). Such small-rated loads 544 may beassociated with in-flight entertainment systems, interior and exteriorlighting systems, sensor heaters, and the like.

In a further aspect of this embodiment, the second bus 515 b can supplyelectrical power to a plurality of adjustable speed motors 552 on theaircraft. Such motors can include cabin pressurization compressors,environmental control system fans, vapor or air cycle ECS packs, largehydraulic pumps, flight actuators, and the like. Use of a high voltageDC system can avoid potential harmonic distortion problems oftenassociated with motor controllers, and can provide a means foraccommodating re-generative energy often associated withelectro-hydrostatic actuators. In addition, the use of a high voltage DCsystem can also provide a significant weight savings through utilizationof lightweight DC generators and the elimination of harmonic distortiontreatment devices and regenerative energy absorption devices.

FIG. 6 is a schematic diagram of an aircraft electric power distributionsystem 600 configured in accordance with another embodiment of theinvention. In one aspect of this embodiment, the power distributionsystem 600 includes a first aircraft engine 610 a, a second aircraftengine 610 b, and an APU 630. The power distribution system 600 canfurther include three AC generators 616 (identified individually as afirst AC generator 616 a, a second AC generator 616 b, and a third ACgenerator 616 c), and three DC generators 618 (identified individuallyas a first DC generator 618 a, a second DC generator 618 b and a thirdDC generator 618 c). The first AC generator 616 a and the first DCgenerator 618 a can be operably coupled to the first engine 610 a.Similarly, the second AC generator 616 b and the second DC generator 618b can be operably coupled to the second engine 610 b. The third ACgenerator 616 c and the third DC generator (318 c can be operablycoupled to the APU 630. The third AC generator 616 c can provideelectrical power from the APU 630 to two AC buses 615 a to service ACloads (not shown) during ground operations and flight as needed. Thethird DC generator 618 c operably coupled to the APU 630 can provideelectrical power to two DC buses 615 b to service adjustable speedmotors (also not shown) during ground operations and flight as needed.In addition, the third AC generator 616 c Can also provide power to thetwo DC buses 615 b via an AC-to-DC conversion device 624. Each of thetwo DC buses 615 b can be operably connected to a corresponding motorcontroller 660 (identified individually as a first motor controller 660a and a second motor controller 660 b). The motor controllers 660 can beconfigured to selectively provide electrical power to either cabinpressurization compressors 680 (identified individually as a firstcompressor 680 a and a second compressor 680 b) or engine start circuits662 (identified individually as a first start circuit 662 a and a secondstart circuit 662 b). In another aspect of this embodiment, the powerdistribution system 600 can include a first electrical receptacle 613 aand a second electrical receptacle 613 b configured to receive powerfrom external ground power sources. In one embodiment, the firstreceptacle 613 a can be configured to receive 115 Vac power from aground source and the second receptacle 613 b can be configured toreceive 230 Vac power from an external ground source.

In one embodiment, high voltage (e.g., 230 Vac) ground power receivedthrough the second receptacle 613 b can be used to start the engines610. In this embodiment, the motor controllers 660 are switched so thatpower from the DC buses 61 5 b is directed to the corresponding enginestart circuit 662. This power is directed to the corresponding ACgenerator 616 and used to run the AC generator 616 as a synchronousmotor to crank the corresponding engine 610 for starting. Once theengine 610 is started, the motor controller 660 switches back to provideelectrical power to the cabin pressurization compressor 680.

FIG. 7 illustrates a schematic diagram of an electric power distributionsystem 700 having only AC generators in accordance with anotherembodiment of the invention. The power distribution system 700 includesa first engine 710 a, a second engine 710 b, and an APU 730. In oneaspect of this embodiment, first and second AC generators 716 a and 716b are operably coupled to the first engine 710 a, third and fourth ACgenerators 716 c and 716 d are operably coupled to the second engine 710b, and fifth and sixth AC generators 716 e and 716 f are operablycoupled to the APU 730. To meet DC voltage needs, high voltage AC powerfrom the engines 710 and the APU 730 can be converted to high voltage DCpower by one or more AC-to-DC conversion devices, such as autotransformer rectifier units (ATRUs) 724, that receive AC power from ACbuses 715. Use of the ATRUs 724 allows the power distribution system 700to provide both high voltage AC and DC power to support conventional 115Vac and 28 Vdc bus architectures. In addition, AC power from one or moreof the AC generators 716 can also be converted to DC power for a 28 Vdcbus 719 by unregulated transformer rectifier units 725 and regulatedtransformer rectifier units 726. In another aspect of this embodiment,having two AC generators 716 coupled to each of the engines 710 allowsboth of the AC generators 716 to be used as synchronous starting motorsfor added engine starting power, if desired. As discussed above withreference to FIG. 3, the embodiments of the invention described abovewith reference to FIGS. 5-7 are not limited to the particular number ofengines and/or starter/generators illustrated, but extend to otherquantities of engines and starter/generators in differentconfigurations.

FIGS. 8A-8C are schematic diagrams illustrating an electric powerdistribution system 800 having an engine start circuit configured inaccordance with an embodiment of the invention. Referring first to FIG.8A, in one aspect of this embodiment, the power distribution system 800includes a generator 816 operably coupled to an engine 810, and acompressor motor 880 operably coupled to an environmental control system840. The generator 816 can provide electric power to a motor controller860 via an AC bus 815 a, an AC-to-DC conversion device 824, and a highvoltage DC bus 815 b. During normal operation as illustrated in FIG. 8A,the motor controller 860 can selectively direct the electric power tothe compressor motor 880 for operation of the ECS 840.

FIG. 8B illustrates an engine starting configuration of the powerdistribution system 800. Because the engine 810 is initially not runningin this configuration, power is provided to the motor controller 860 byan alternate AC power source 830 rather than the generator 816. In oneembodiment, the alternate power source 830 can include an APU or anexternal power source. In one aspect of this embodiment, the motorcontroller 860 selectively directs the electric power from the alternatepower source 830 to an engine start circuit 862. The engine startcircuit 862 provides the electrical power to the generator 816, which isconfigured to run as a synchronous motor for starting the engine 810.

FIG. 8C illustrates another engine starting configuration of the powerdistribution system 800. Here, power for starting the engine 810 isprovided by a battery 836. In one aspect of this embodiment, the motorcontroller 860 selectively opens the circuit to the high voltage DC bus815 b so it can receive electric power from the battery 836. Afterconnecting to the battery 836, the motor controller directs the electricpower to the generator 816 via the engine start circuit 862 as describedabove.

From the foregoing, it will be appreciated that specific embodiments ofthe invention have been described herein for purposes of illustration,but that various modifications may be made without deviating from thespirit and scope of the invention. Accordingly, the invention is notlimited, except as by the appended claims.

1. An aircraft comprising: a fuselage; a jet engine configured toprovide propulsive thrust to the aircraft; an electric generatoroperably coupled to the jet engine and configured to receive shaft powerfrom the engine; an environmental control system configured to provideconditioned air to at least a portion of the fuselage in the absence ofbleed air from the jet engine, the environmental control systemincluding at least one fan motor configured to receive electric powerfrom the electric generator; and a hydraulically actuated landing gearconfigured to movably support at least a portion of the aircraft on theground, the landing gear receiving hydraulic power from a hydraulic pumpdriven by an electric motor in the absence of bleed air from the jetengine, the electric motor receiving electric power from the electricgenerator.
 2. The aircraft of claim 1 wherein the electric generatoroperably coupled to the jet engine is a DC generator, and wherein theaircraft further comprises: an AC generator operably coupled to the jetengine and configured to receive shaft power from the jet engine; a wingextending outwardly from the fuselage; and an electrothermal wing iceprotection system configured to at least reduce the formation of ice ona portion of the wing, the electrothermal wing ice protection systemconfigured to receive electric power from the AC generator in theabsence of bleed air from the jet engine.
 3. The aircraft of claim 1wherein the electric generator operably coupled to the jet engine is aDC generator, wherein the aircraft further comprises an AC generatoroperably coupled to the jet engine and configured to receive shaft powerfrom the jet engine, and further wherein the AC generator is operable asa synchronous motor to start the jet engine in the absence of pneumaticpower.
 4. The aircraft of claim 1 wherein the electric generatoroperably coupled to the jet engine is a DC generator, and wherein theaircraft further comprises: an AC generator operably coupled to the jetengine and configured to receive shaft power from the jet engine; and anAC-to-DC conversion device configured to receive AC power from the ACgenerator and direct DC power to the fan motor of the environmentalcontrol system.
 5. The aircraft of claim 1 wherein the electricgenerator operably coupled to the jet engine is a DC generator, andwherein the aircraft further comprises: an AC generator operably coupledto the jet engine and configured to receive shaft power from the jetengine; an AC-to-DC conversion device configured to receive AC powerfrom the AC generator; and a motor controller configured to receive DCpower from the AC-to-DC conversion device and direct the DC power to thefan motor of the environmental control system, wherein the motorcontroller is further configured to selectively direct DC power from asource other than the AC generator to the AC generator to operate the ACgenerator as a synchronous motor to start the jet engine in the absenceof pneumatic power.
 6. On an aircraft having a fuselage and a jet engineconfigured to provide propulsive thrust, a system for providingconditioned air to the fuselage, the system comprising: means forproviding outside air to a compressor fan in flow communication with thefuselage; means for extracting electric power from the jet engine; andmeans for providing at least a portion of the electric power to thecompressor fan to flow the outside air from the compressor fan into thefuselage of the aircraft in the absence of bleed air from the jetengine.
 7. The system of claim 6, further comprising means for providingat least a portion of the electric power from the electric generator toa heating element of a wing ice protection system, the wing iceprotection system configured to at least reduce the formation of ice ona portion of the wing in the absence of bleed air from the jet engine.8. The system of claim 6, further comprising means for providing atleast a portion of the electric power from the electric generator to anelectric motor-driven hydraulic pump of a landing gear system, thehydraulic pump configured to operate a landing gear in the absence ofbleed air from the jet engine.
 9. The aircraft of claim 1, furthercomprising a motor controller operably coupled to the electric generatorto vary the electric power provided by the electric generator to theenvironmental control system based on the power demands of the fanmotor.
 10. The system of claim 6, further comprising means for varyingthe electric power provided to the compressor fan by the jet enginebased on the power demands of the compressor fan.
 11. An aircraftcomprising: a fuselage having a passenger cabin; a jet engine configuredto provide propulsive thrust to the aircraft; an electric generatoroperably coupled to the jet engine and configured to receive shaft powerfrom the jet engine; an environmental control system including at leastone compressor motor configured to receive electric power from theelectric generator to provide outside air to the passenger cabin in theabsence of bleed air from the jet engine; and a motor controlleroperably coupled to the electric generator to vary the electric powerprovided by the electric generator to the compressor motor based on thepower demands of the environmental control system.
 12. The aircraft ofclaim 11, further comprising: a wing extending outwardly from thefuselage; and an electrothermal wing ice protection system configured toat least reduce the formation of ice on a portion of the wing in theabsence of bleed air from the jet engine, the electrothermal wing iceprotection system configured to receive variable electric power from theelectric generator.
 13. The aircraft of claim 11, further comprising: awing extending outwardly from the fuselage; and an electromechanicalwing ice protection system configured to at least reduce the formationof ice on a portion of the wing in the absence of bleed air from the jetengine, the electromechanical wing ice protection system configured toreceive variable electric power from the electric generator.
 14. Theaircraft of claim 11, further comprising: a wing extending outwardlyfrom the fuselage; and a wing ice protection system configured to atleast reduce the formation of ice on a portion of the wing in theabsence of bleed air from the jet engine, the wing ice protection systemconfigured to receive electric power from the electric generator in acycled manner.
 15. The aircraft of claim 11, further comprising: ahydraulically actuated landing gear extendable downwardly from theaircraft; a hydraulic pump configured to provide hydraulic power to thelanding gear; and an electric motor operably coupled to the hydraulicpump and configured to receive electric power from the electricgenerator to drive the hydraulic pump in the absence of pneumatic powerfrom the jet engine.
 16. The aircraft of claim 11 wherein the electricgenerator is a starter/generator operable as a synchronous motor tostart the jet engine in the absence of pneumatic power.
 17. The aircraftof claim 11 wherein the electric generator is a starter/generatoroperable as a synchronous motor to start the jet engine, and wherein thejet engine is configured to be started by the starter/generator in theabsence of a pneumatically operable starter turbine.
 18. The aircraft ofclaim 11 wherein the electric generator operably coupled to the jetengine is a first electric generator, and wherein the aircraft furthercomprises: an auxiliary power unit; and a second electric generatoroperably coupled to the auxiliary power unit and configured to receiveshaft power from the auxiliary power unit, wherein the at least onecompressor motor of the environmental control system is configured toreceive electric power from the second electric generator to provideoutside air to the passenger cabin in the absence of compressed air fromthe auxiliary power unit.
 19. The aircraft of claim 11 wherein thecompressor motor of the environmental control system is an adjustablespeed motor configured to vary compressor speed in response to changesin pressurization demands of the fuselage.
 20. The aircraft of claim 11,further comprising: a fuel tank; and a variable-speed fuel pumpconfigured to transfer fuel from the fuel tank to the jet engine atvariable speeds based on the demand for fuel by the jet engine, whereinthe fuel pump is configured to receive electric power from the electricgenerator.